A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine general includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
The combustion section typically includes a plurality of individual fuel nozzles. The plurality of individual fuel nozzles are circumferentially spaced and configured for injecting a fuel-air mixture to be burned within a combustion chamber. Although designed to create a substantially homogenous mixture of combustion gasses along the circumferential direction, typically, a location downstream of each fuel nozzle is hotter than other circumferential positions (sometimes referred to as “hot gas streaks”).
Accordingly, when such flow is routed into the turbine section, the hot gas streaks may be hotter than other circumferential locations. In order to ensure the components within the turbine section are capable of withstanding the temperatures of the hot gas streaks, each of the components within the turbine section are designed and manufactured to accommodate these hot gas streaks. However, the inventor of the present disclosure has discovered that such an approach may lead to over engineering of certain components within the turbine section.
Accordingly, the inventor of the present disclosure has discovered that it may be beneficial for components within the gas turbine engine to be designed according to their anticipated or actual location within the turbine section relative to the hot gas streaks extending therethrough during operation. More specifically, the inventor of the present disclosure has discovered that it may be beneficial to design components within a given stage of turbine components according to their anticipated or actual location relative to the hot gas streaks extending through the turbine section during operation.